Typically, gas turbine engines include a compressor for compressing air, a combustor for mixing the compressed air with fuel and igniting the mixture, and a turbine blade assembly for producing power. Combustors often operate at high temperatures that may exceed 2,500 degrees Fahrenheit. Typical turbine combustor configurations expose turbine blade assemblies to these high temperatures. As a result, turbine blades must be made of materials capable of withstanding such high temperatures. In addition, turbine blades often contain cooling systems for prolonging the life of the blades and reducing the likelihood of failure as a result of excessive temperatures, particularly in concentrated areas of over temperature, sometimes referred to as hot spots.
Typically, turbine blades are formed from a root portion having a platform at one end and an elongated portion forming a blade that extends outwardly from the platform coupled to the root portion. The blade is ordinarily composed of a tip opposite the root section, a leading edge, and a trailing edge. The inner aspects of most turbine blades typically contain an intricate maze of cooling channels forming a cooling system. The cooling channels in a blade receive air from the compressor of the turbine engine and pass the air through the blade. The cooling channels often include multiple flow paths that are designed to maintain all aspects of the turbine blade at a relatively uniform temperature.
Size and space limitations make trailing edges one of the more challenging sections of a turbine blade to cool. Traditionally, a trailing edge camber line discharge together with pin fins or multiple impingements has been used airfoil trailing edge region cooling. Such design requires a thicker trailing edge that can induce higher aerodynamic blockage and reduce stage performance. Techniques for cooling a thinner trailing edge have been developed. For example, a first stage blade can utilize a pressure side bleed the exhausts on the pressure side adjacent to the tip of the trailing edge, rather than a camber line discharge at the center of the trailing edge. This cooling channel arrangement allows for a reduction in the effective thickness of the trailing edge when compared to the required thicknesses of both the suction side and pressure side regions of the trailing edge surrounding a camber line cooling discharge.
However, the pressure side bleed cooling approach causes a side flow and presents shear mixing between the cooling air and the mainstream flow as the cooling air exits the pressure side channel outlet. The shear mixing of the cooling air with the mainstream flow reduces cooling effectiveness of the trailing edge overhang, thus inducing over temperature or a hot spot at the trailing edge suction side location. Frequently, a hot spot can become the life limiting location for the entire airfoil. Thus, a need exists for a cooling system capable of providing sufficient cooling to a relatively thinner trailing edge of a turbine airfoil.